The present disclosure relates generally to turbine engines and, more specifically, systems and methods of cooling compressor bleed air.
At least some known gas turbine engines include a fan, a core engine, and a power turbine. The core engine includes at least one compressor, a combustor, and a high-pressure turbine coupled together in a serial flow relationship. More specifically, the compressor and high-pressure turbine are coupled through a shaft to form a high-pressure rotor assembly. Air entering the core engine is mixed with fuel and ignited to form a high energy gas stream. The high energy gas stream flows through the high-pressure turbine to rotatably drive the high-pressure turbine such that the shaft rotatably drives the compressor. After being discharged from the high-pressure turbine, the gas stream continues to expand as it flows through a low-pressure turbine positioned aft of the high-pressure turbine. The low-pressure turbine includes a rotor assembly coupled to a drive shaft and a fan. The low-pressure turbine rotatably drives the fan through the drive shaft.
Many modern commercial turbine engines operate at increasingly high temperatures to facilitate increasing engine performance and efficiency. However, operating at the increasingly high temperatures can result in damage to hot gas path components over extended periods of operation. As such, at least some known turbine engines include an active cooling system that facilitates reducing a temperature of hot gas path components. For example, the hot gas path components are typically cooled with a flow of compressor bleed air discharged from the high-pressure compressor. In some applications, the compressor bleed air is cooled with a flow of bypass air channeled from a fan duct. However, bleeding bypass air from the fan duct can result in an increase in drag on the turbine engine assembly during flight.